As one skilled in this art appreciates, the compressor in a gas turbine engine is to provide high pressure air which is heated in the combustion section and expands over the turbine where energy is extracted to power the compressor and provide thrust for propelling the aircraft. The amount of energy left over after powering the compressor represents the useful jet thrust of the engine and this excess energy is very sensitive to the compressor efficiency. It is therefore imperative that the compressor operate at a high efficiency. Hence, the engine is designed so that the compressor pressurizes the most air through the particular configuration of the flow path in the compressor section consistent with optimum compressor efficiency over the range of engine operating conditions.
The operation of the compressor depends primarily on the lift force of the airfoils of the compressor blades. The approaching air flowing over these blades exerts lift and drag forces. If the angle of attack of the approaching air stream becomes too high the air resulting in an increase in drag and decrease in lift. If the angle of attack varies too far in the other direction the airflow will separate from the surface of the blade again increasing drag. If the speed of the airflow exceeds a certain value, the approach Mach number becomes too high and the airflow accelerating to pass around the airfoil will exceed the speed of sound and a shock wave will result cause turbulent flow and again an increase in drag.
In designing the compressor, the designer not only takes into consideration the flow velocity and the rotative speed of the compressor, but also must consider the cascade effect in a multiple stage of an axial compressor. In addition to being compatible with the remaining portion of the engine, namely, the combustor and turbine, consideration must be given to the mechanical aspect of the compressor. Once the inlet hub tip ratio of the compressor is determined which at best is a compromise between specific airflow and weight, the number of stages must be determined. Once the slope of the compressor is determined and the blade velocities and air velocities are known the number of stages can be determined. The work at each stage of compression will be dictated in accordance with the Mach number and stalling limits of the cascade.
Even after selecting the slope of the compressor which can be either the inner or other diameter, the number of stages and the velocities of the rotation of the blades and the velocity of the airflow, under certain operating conditions of an advanced gas turbine engine it was found that the compressor became highly loaded. It was therefore a problem to keep the stators from becoming overloaded resulting in a detriment to the efficiency of the compressor.
While traditionally the compressor end wall flow path is designed so that the rotor and stator leading and trailing edge diameters can be described by a smooth spline, we have found that we can solve the problem of overloading the stator by increasing the amount of flow path convergence across the stator. In accordance with this invention, we made the annulus area between blade rows constant. This serves to eliminate re-acceleration of the flow field and consequently reduce loss in static pressure.